In-space propulsion is an important part of most space missions, accounting for an average of 40% of the launched mass of unmanned systems. Pulsed bipropellant thrusters are commonly used on spacecraft for in-space propulsion. Short-pulse operation is particularly useful for attitude control, the figure of merit being the minimum impulse bit. Propellant consumption in limit cycle operations is particularly sensitive to the minimum impulse bit. A rough estimate of the size of the minimum impulse bit of a thruster is the steady state force generated times the pulse width, but significant deviations from this concept are found experimentally. Pulse widths as short as 0.01 s are used operationally. The efficiency of pulsed bipropellants has been estimated as 50% of the theoretical limit for 0.01 s pulses, rising to 75%-85% for 0.1 s pulses. This can be compared to a typical steady state efficiency of 92% of theoretical. In addition to the reduced efficiency of pulsed mode operation, there is also an increased variability. The thrust can vary by 10% from pulse to pulse, or ±30% from engine to engine.
Incomplete combustion of pulsed bipropellant thrusters and can lead to condensed phase reaction intermediates that can contaminate spacecraft surfaces, and even cause erosion. These contaminants are usually seen on start-up or shutdown, and there is a correlation between the amount of contaminant formed, and the amount the specific impulse of the pulsed mode is reduced. In some cases, the reaction intermediates formed in incomplete combustion can lead to hard-starts, or engine damage.
There are several reasons for the differences between steady state and pulsed operation. The start-up of a thruster on-orbit is significantly different from the steady state operation. The combustion chamber is under vacuum and cold, and the propellant lines from the thruster valve through the injector are also under vacuum. Just the difference in temperature is enough to cause significant differences in the flow field. Other differences compound the effect. The time required for ignition, and the establishment of steady state combustion chamber pressure and temperature is relatively short compared to typical thruster burns. For a short pulse meant to deliver the minimum impulse bit, however, the combustion chamber may never reach steady state conditions within the pulse. Incomplete combustion could be one of the factors contributing to the low efficiency and high variability of pulsed thrusters, as well as the generation of contaminants. An additional potential concern is the freezing of propellant on exposure to space vacuum.
The hydraulic diameter of dribble channels and injectors affect propellant flow properties of injector channels of bipropellant thrusters. The dribble channels are used to synchronize the response times of the fuel and oxidizer. The synchronization of the propellant response times allows the mixture ratio to be controlled over all burn times. Bipropellant engines are used for multiple purposes on spacecraft. These purposes include orbit insertion, orbit trim, plane changes, repositioning, station keeping, attitude control, maneuvering, reaction wheel unloading, and thrust vector control. Pulsed bipropellant engines are most often used in attitude control. The performance of pulsed bipropellant engines is judged by the specific impulse, ISP, and the minimum impulse bit, IMIN. In limit cycle operations, for example, where a pair of thrusters are used to keep the attitude of the spacecraft within a set angular range, the propellant usage WP is very sensitive to IMIN. A WP usage equation is defined by WP=n2(IMIN)2L/4ISPIVθL, where n is the number of motors firing, L is the radius from the vehicle center of mass to the thrust vector, IV is the mass moment of inertia of the vehicle, θL is the angular limit, or one half the allowed coasting angle between firings. A smaller IMIN could reduce propellant usage or alternatively be used to provide tighter angular control at the same propellant consumption rate. IMIN is set by the minimum pulse width ISP, and the way the engine responds to the short pulse of propellant. Current bipropellant thrusters have an IMIN of 0.09 N-s for a 20.0 ms pulse. The ISP in short 20.0 ms pulses is about half that at steady state. 100 ms pulses have about three-quarters the ISP of steady state firings. The variability of IMIN is about 10% from pulse to pulse and 30% from engine to engine.
Monopropellant engines have lower IMIN, but also a lower ISP. For missions where the required total impulse is less than 45,000 N-s monopropellant systems are used because the propellant savings does not justify the increased complexity. Monopropellant systems usually use hydrazine, whereas bipropellant systems usually use monomethyl hydrazine and nitrogen tetroxide. An alternate design is the dual mode system, which uses hydrazine and nitrogen tetroxide in a bipropellant engine for orbit insertion and other large velocity change propulsion requirements. The hydrazine in these systems can then be used for monopropellant engines for attitude control. The use of these systems highlights the need for a good, pulsed bipropellant engine.
Standard bipropellant injector design uses a single valve to control the flow of both the fuel and oxidizer. The propellant line between the thruster valve and the injector orifice is called the dribble volume. The dribble volume in a dribble channel is kept small to minimize start up time, but needs to be large enough to limit soak back of heat from the combustion chamber. The dribble volume consists of a feed tube, that is, a dribble channel, and an injector orifice, or simply, an injector. The injector orifice is a smaller diameter hole or tube terminating on the injector face. Fuel and oxidizer are injected into the combustion chamber through an equal number of circular orifices. The injector orifices are sized to achieve the desired mass flow ratio of oxidizer to fuel, the mixture ratio, for steady state operation. To facilitate start up of the thruster, the oxidizer dribble volume is smaller, allowing the oxidizer to reach the combustion chamber first.
Discovery by the present inventor is made of an unknown problem that an initial propellant flow or the flow during a short pulse of the thruster valve can be much slower than the steady state value. The response time of the propellant flow is limited by inertia and viscous drag. The response time of a particular propellant is given by a response time equation τ=ρα/μ, where τ is the response time, ρ is the density, α the hydraulic radius, and μ is the viscosity. For current injector designs, the response time of the oxidizer flow is three or four times longer than the response time of the fuel flow, and both response times are many times the length of the shortest pulses typically used on orbit. The mismatch in the response times of the propellant flows means that for short pulses, and during the initial phases of steady state operation, the propellant mixture ratio will be extremely fuel rich. The few millisecond oxidizer lead provided by a smaller oxidizer dribble volume in the current designs is not enough to compensate for this mismatch.
The fuel rich mixture results in lower combustion temperatures, and incomplete combustion, which causes several problems. Incomplete combustion can lead to the accumulation of combustion intermediates. These intermediates are often called fuel oxidizer reaction products. The fuel oxidizer reaction products produce undesirable high-energy condensation. The products have been linked to several problems with thrusters, such as hard starts, and even damage to the thruster body. The fuel oxidizer reaction products can also contaminate other spacecraft surfaces where there is a lower vapor pressure than fuel, oxidizer, or combustion products. Low chamber temperatures make fuel oxidizer reaction product accumulation more likely. Incomplete combustion and the accumulation of fuel oxidizer reaction products are responsible for the low ISP and high variability of the IMIN in pulsed operation. The accumulation of fuel oxidizer reaction products and impacts on engine operation have limited the minimum pulse length used in bipropellant engines, and thus the IMIN.
Unequal propellant response times can also lead to combustion instability. Normally, propellant flow is believed to have a damping effect on pressure oscillations. When the pressure in the combustion chamber spikes up due to some perturbation, the propellant flow will decrease due to the reduced pressure drop from the tank to the combustion chamber. The reduced propellant flow will lead to a lower chamber pressure countering the effect of the perturbation. When, however, the fuel and oxidizer flows respond to the pressure fluctuation on different time scales, the pressure spike can be amplified.
Bipropellant engines usually run fuel rich. The rich mixture ratio keeps the combustion chamber temperature lower with the excess fuel being used to cool the chamber walls. The excess fuel lowers the average mass of the exhaust gas for yielding a slightly higher exhaust velocity. When the fuel flow responds more quickly to a positive pressure fluctuation, there is a shift in the mixture ratio closer to stoichiometric. The stoichiometric mixture burns hotter, and causes a larger pressure increase, amplifying the fluctuation. Synchronization of the propellant response times will prevent shifts in the mixture ratio, and lead to damping of pressure fluctuations, and more stable combustion. Prior propellant injector systems have used conventional dribble channels and injectors producing mismatches in fuel and oxidizer delivery, contamination, and spikes in the combustion performance. These and other disadvantages are solved or reduced using the invention.